1. Field of the Invention
The present invention relates to gas turbine combustors and more specifically to a flow sleeve having an inlet region that reduces pressure loss to the compressed air entering a combustor.
2. Description of Related Art
A gas turbine engine typically comprises a multi-stage compressor, which compresses air drawn into the engine to a higher pressure and temperature. A majority of this air passes to the combustors, which mixes the compressed heated air with fuel and contains the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which drives the compressor, before exiting the engine. In land-based gas turbines, the turbine is also coupled to a generator for generating electricity.
For land-based gas turbine engines, often times a plurality of combustors are utilized. Each of the combustion systems include a case that serves as a pressure vessel containing the combustion liner, which is where the high pressure air and gas mix and react to form the hot combustion gases. Typically the case is fabricated from a lower temperature capable material such as carbon-steel. In order to ensure that the case is not overexposed to the temperatures of the combustion liner as well to ensure that the combustion liner receives the proper amount of air for cooling and mixing with the fuel, an additional liner is often located within the case and is coaxial to the combustion liner and case. This additional liner is more commonly referred to as a flow sleeve.
A two-stage combustion system of the prior art commonly used in land-based gas turbine engines is shown in cross section in FIG. 1. Combustor 10 includes a generally annular case 11 having a center axis A—A and an end cover 12 that is fixed to a case flange and contains a plurality of fuel nozzles 13 located about center axis A—A. Located coaxial to center axis A—A is a combustion liner 14 having a first combustion chamber 15 and second combustion chamber 16, separated by venturi 17 having a throat of reduced cross sectional area 18. An additional fuel nozzle 19 is located along center axis A—A. Located coaxial to combustion liner 14 and radially between case 11 and combustion liner 14 is flow sleeve 20. As mentioned previously, flow sleeve 20 serves to direct compressed air along the outer walls of liner 14 for cooling purposes, as well as for being injected to mix with the fuel for combustion. In combustor 10 of the prior art, flow sleeve 20 forms a generally annular passageway 21 around combustion liner 14 for directing the required amount of compressed air to combustion liner 14 for cooling and mixing with the fuel from fuel nozzles 13 and 19. In prior art combustor 10, compressed air is introduced to the combustion system through a generally annular flow sleeve inlet 22, which is shown in a more detailed cross section in FIG. 2.
Flow sleeve inlet 22 is formed between flow sleeve 20 and transition duct 25, which has a bellmouth portion 26 and a structural support ring 27, each of which are located towards the forward end of transition duct 25. In this combustor configuration, bellmouth 26 and support ring 27 create obstructions that block or disturb a portion of the compressed air flow that enters passageway 21 through flow sleeve inlet 22, thereby causing an undesirable pressure loss to the air supply. This disturbance to the air flow and resulting pressure loss has multiple negative effects on the hardware durability and performance. Specifically, hula seal 28, which, in the prior art, is a seal encompassing the aft end outer surface of liner 14 and contains a plurality of axial slots that form “fingers” that spring to seal between liner 14 and transition duct 25, does not receive sufficient cooling air due to a separation zone 29 created by air flow passing over bellmouth 26 (see FIG. 2). As a result of this lack of cooling air, the aft end of combustion liner 14 and hula seal 28 operate at a higher temperature, causing more radial interference between hula seal 28 and transition duct 25 than desired, leading to premature wear of hula seal 28. The flow disturbances created by bellmouth 26 and ring 27 combined with the geometry of flow sleeve inlet 22, due to the axial length of the aft region of flow sleeve 20, creates a pressure loss to the incoming air supply. The pressure loss at flow sleeve inlet 22, which is approximately 1.5% of the available air pressure, results in a lower cooling air supply pressure to combustion liner 14. Annular passageway 21 creates little, if any, additional pressure loss to the cooling air. As a result, less air is passed through the various passages requiring cooling and injected for mixing with the fuel, thereby resulting in higher operating temperatures, a less durable design, and reduced combustor performance. As one skilled in the art of gas turbine combustion will understand, maintaining adequate cooling of the combustion liner is imperative for combustor durability and performance.
Therefore, what is needed is a flow sleeve for a gas turbine combustor having an inlet region that reduces the pressure loss to the incoming compressed air, such that a high enough air pressure is available to provide sufficient cooling to the combustion liner surfaces. This is especially true for combustors that operate for an extended period of time and require large amounts of cooling and enhanced mixing in order to achieve low emissions.